1. Field of the Invention
The present invention relates generally to aircraft mounted gas turbine engines, and, more specifically, to a thrust mount which substantially eliminates engine case deflection caused by engine backbone bending due to thrust loads, and which also minimizes casing ovalization.
2. Description of Related Art
A principal type of modern aircraft gas turbine engine in usage today is of the turbofan type. All of the working medium gases are directed through the fan section of the engine. A portion of the working medium gases from the fan section exit is directed through the compression, combustion, and turbine sections of the engine core. The remaining fan exhaust is directed outboard around the engine core. Both the fan discharge and the core discharge flows produce thrust and can be mixed together prior to discharge to improve thermal efficiency or can be discharged individually. The diameter of the engine at the fan stages is typically quite large, on the order of ten feet and larger for high bypass ratio engines of the eighty thousand pound thrust class. Another type of aircraft gas turbine engine is the turboprop engine where a core engine drives unducted props through a gearbox and/or with a free turbine operating with the hot turbine exhaust gases from the core engine or gas generator as it is often referred to. Another type of engine presently under development is the very high bypass ratio engines of both the ducted and unducted type using both pusher and puller fans and which are often driven by a free turbine which is powered by a gas generator or core engine.
Each engine is supported by an aircraft structure, for example, on a pylon extending downwardly beneath the wing or, as in the case of the present invention, sideways on a pylon extending horizontally or sideways from the aircraft fuselage usually at or near the end and tail of the aircraft. The engine is typically mounted and secured to the aircraft by two connections, one towards the forward end of the engine, usually just rearward of the fan section and a second toward the aft end of the engine, typically in the turbine section. The engine static structure is made up of casings and frames. The engine mount is usually attached to an engine frame, a static structure, which supports the rotating components generally referred to as rotors. The engine static structure generally has sub-structures including a forward frame and an aft frame connected by a core engine casing often referred to as a backbone. Forward and aft frames having radially extending structural struts typically support the engine bearings which in turn rotatably support the rotors within hubs of the frames. Typically a dual rotor engine has a forward fan frame and a rear turbine frame that support the main rotor bearings. Many engines have intermediate frames such as an mid-turbine frame. Frames can support more than one rotor.
The engine casing usually is suspended from the pylon by a forward mount assembly that extends horizontally from the aircraft to interconnect with a forward portion of the engine casing, such as the fan casing, and an aft mount assembly that extends horizontally from the aircraft to interconnect a rearward section of the engine core, such as a turbine frame. Thrust produced by the engine are reacted out through the engine and to the aircraft by the mount assemblies and separate engine thrust mounts.
The advent of large, high bypass turbo-fan jet propulsion engines with their greater flexibility has resulted in relatively large deflections occurring between the engine casings and the rotors of engines that have been mounted to the aircraft by conventional means. This results in rubbing contact between the rotor blade tips and the engine casings or engine designs with larger than desired tip clearances between the rotor and the stator to avoid rubs between the rotor and the stator. When the engines are operated at full power, such as during takeoff, the high thrust loads that act through the engine must be reacted by the engine thrust mounts. Since the engine thrust mounts are generally offset from the longitudinal center of the engine along and through which the thrust acts, bending moments are generated in the engine cases by the offset. The large bending loads resulting therefrom cause deflection of the engine components resulting in interference between the rotor blades and seals and their associated casings.
Excessive blade and seal wear increases the clearance between these components causing a loss of fuel efficiency. In addition, a phenomenon known as blade tip stall may result from larger clearances between the blade tips and the engine casings. This can lead to vibrational problems associated with non-synchronous whirl motion of the rotors. Moreover, when compressor blades rub against their surrounding seal, particles are removed from the blade tips and the seal. The deposition of these particles on the extremely hot turbine sections of the engine can plug cooling holes and roughens the turbine blades and stators and reduces their aerodynamic efficiency. Interstage seals can also wear which will open clearances and reduce fuel efficiency.
One standard technique for eliminating tip rub has been simply to provide larger clearances between the blades and the casing or selectively pre-grind the sections of the casing which are susceptible to tip rub. Typically, clearances are set to be minimum at maximum thrust with the objective of being as small as possible at cruise where fuel consumption has a big impact on aircraft system efficiency. Although this may assist in avoiding blade and seal particles from being deposited on the hot engine sections and avoiding loss of blade tip material, it results in lower fuel efficiency and can cause detrimental vibration of the rotors.
Another approach to reducing tip rubs is to thicken the engine casings and increase the number of rotor bearings to provide stiffening and better concentricity between rotors and stationary parts (casings, seals, frames). This approach, however, is very costly in terms of weight, complexity, and cost. One proposed design, disclosed in U.S. Pat. No. 4,022,018, provides a jet engine mounted in a nacelle structure having a hydraulic actuator disposed between the bottom of the fan casing and the nacelle structure in an attempt to restrain the fan casing against movement relative to the nacelle structure.
U.S. Pat. No. 4,326,682 provides a system for mounting a jet engine sideways to a boom with forward and aft linkages used to fixedly suspend the fan and turbine casing of a turbo fan engine to the boom. The mounts are laterally extending links attached to the casings and boom by clevises which fixedly hold the engine in an axial position with respect to the boom. A thrust tie extends from the rear mount to the inner casing of the engine at an oblique angle such that it intersects the engine centerline in the plane of the front links. The problem with such a design, besides being limited to sideways mountings, is that thrust loads are taken out through outer casings and frames as well as through the thrust ties.
U.S. patent application Ser. No. 07/857,136, now U.S. Pat. No. 5,320,307, entitled "Aircraft Engine Thrust Mount" filed Mar. 25, 1992, and incorporated herein by reference, discloses a gas turbine engine which is typically mounted below an aircraft wing to a pylon at its forward end, an intermediate section, and its aft end for transmitting loads to the pylon. The loads typically include vertical loads such as the weight of the engine itself, axial loads due to the thrust generated by the engine, side loads such as those due to wind buffeting and side accelerations, and roll loads or torques due to rotary operation of the engine. The forward and aft vertical mounts are so designated because they typically carry in-plane loads to the pylon wherein the in-plane loads are those occurring in a single axial plane extending perpendicularly outwardly from the longitudinal axis or centerline of the engine and include vertical and horizontal loads and rotary torque or moments. An aft thrust mount is provided for transferring the axially directed thrust loads from the engine to the pylon which are tension loads during forward propulsion of the aircraft, and compression loads which occur during the use of the engine's thrust reverser during braking of the aircraft upon landing. The thrust mount can be in the front also, in which case the axially directed thrust loads are transferred in tension during thrust reversal and compression during normal thrust.
This thrust mount includes a pair of 90.degree. circumferentially spaced apart elongated thrust links pivotally joined at forward ends thereof to a conventional fan frame and at opposite aft ends pivotally joined to a lever sometimes referred to as a whiffle tree which serves as a mount platform. The two thrust links are pivotally joined to opposite ends of the lever, and the center of the lever is pivotally joined to a platform which is fixedly joined to the pylon. The several thrust link pivotal joints include conventional spherical bearings, which allow slight rotation of the thrust links in three orthogonal planes relative to the fan frame and the lever. And, the lever center joint includes a pin through a bushing for single plane rotation.
It is desirable to provide a load path from the fan frame to the pylon which substantially eliminates bending loads in the engine casing or backbone and the resultant rubbing effects on the rotor, particularly at high power thrust levels. As the size of the thrust load developed by modern turbofan engines has increased, so has the magnitude of the reaction loads and bending moment. An inefficient method of increasing stiffness is to increase part thickness which is a very heavy and costly solution to the problem. It is therefore highly desirable to further reduce the amount of case deflection or backbone bending than that afforded by the elongated thrust yoke mount system in U.S. Pat. No. 4,603,821 and others like it. The resultant engine static structure deflection causes increased rubbing between the rotating hardware and the adjacent stationary hardware. This abrasive wear results in an adverse impact on engine performance and specific fuel consumption, and necessitates more frequent engine maintenance and overhaul. Repair and replacement of rotor blades is one of the highest operating costs for an aircraft gas turbine engine.
Increasing fuel costs and demands for improved durability accentuate the need for low weight designs and systems for substantially eliminating engine case deflection and not just diminishing backbone deflection particularly during high power settings such as at takeoff. The problem is greater for very high bypass ratio engines having both ducted and unducted fans with fan diameters much greater than their core engine diameters. Increasingly, jet engine manufacturers are designing and building engines with increased fan bypass ratios because such engines provide greater propulsive efficiency. There is a great need to substantially eliminate engine case deflection caused by engine backbone bending due to thrust loads.